Turbine nozzle and rotor cooling systems

ABSTRACT

A gas turbine engine having a turbine with internally cooled, stationary nozzles, a rotor with internally cooled blades, and systems for supplying air to the nozzles and rotor and for cooling the nozzle case.

The present invention relates to gas turbine engines and, morespecifically, to novel, improved gas turbine engines having turbinenozzle, nozzle case, and rotor cooling systems.

One of the important characteristics of the turbine engines with whichthe novel cooling systems disclosed herein are particularly useful ishigh turbine inlet temperatures (in one case 1825° F./1014° C.). Suchengines are capable of obtaining high thermal efficiencies withoutrecuperators which is advantageous from the viewpoints of simplicity andreduced cost, weight, and bulk.

However, at temperatures of the magnitude identified in the precedingparagraph, the creep and rupture strengths of available alloys arereduced to levels which are sufficiently low that service lifes of thelength we contemplate cannot be obtained.

The novel cooling systems disclosed herein are capable of limiting thetemperatures of turbine components to a maximum of 1500° F. at turbineinlet gas temperatures in the 1800° F. range. At the lower temperature,30,000 hours (ca. 3.4 years) of continuous duty operation and usefulcomponent lifes of 30 years are attainable.

One turbine engine in which the novel cooling systems disclosed hereinmay be employed to particular advantage is of two-shaft design. It has atwo-stage gas producer turbine for driving the compressor of the engineand a power turbine for driving, as examples, an electrical generator ora booster compressor. The power turbine can also be employed as amechanical drive prime mover.

A first of the novel cooling systems disclosed herein is designed tocool the first and second stages nozzles and the nozzle case of theengine described above.

In this system air is bled from the discharge side of the compressorthrough a longitudinally extending annular passage to a plenum orchamber between the turbine nozzle case and the external turbinehousing. One stream of the compressor discharge air flows through ascreen into the first stage nozzles while a second stream is filteredthrough the screen, impinged against the nozzle case to cool it, andthen directed into the second stage nozzles.

The nozzles in both the first and second stages are cooled by anoptimized combination of impingement, convective, and film cooling. Inboth stages the air flows into a sheet metal insert inside each nozzle.Part of the air exits through holes in the leading edge of the insertproviding impingement cooling of the leading edge of the nozzle. The airthen flows toward the trailing edge of the nozzle through the spacebetween the latter and the insert, convectively cooling the forward partof the nozzle. This air is discharged through slots in the nozzle toproduce film cooling of its aft or trailing edge portion.

The first-stage nozzles have rows of air discharge holes on both thepressure and suction sides of the nozzles. The second stage nozzles havea single row of holes on the pressure side near the trailing edge.

The remainder of the air introduced into the nozzle inserts of both thefirst and second stages exits through holes in the trailing edges of theinserts and flows through the nozzles to convectively cool the aft partsof the nozzles. This air is discharged through slots in the trailingedges of the first stage nozzles and, in the second stage nozzles,through the same slots as the air flowing around the insert from theholes in its leading edge.

In addition there are passages or openings through the inner ends of thesecond stage nozzles. Air discharged through these openings cools asecond stage diaphragm extending between the second stage nozzles andthe first stage of the gas producer turbine rotor.

Also, part of the first cooling air stream flowing through the turbinecase is diverted from the first stage nozzles through the trailing edgesof the segments in which the nozzles are incorporated. This air flowsacross and cools the tip shoes surrounding the first stage of the gasproducer turbine rotor.

A second of the novel systems disclosed herein is employed toconvectively cool the blades of the gas producer turbine first stage,the first stage rotor disc, and the first and second stage nozzlediaphragms.

Each of the first stage blades has a three-pass internal coolingcircuit. Compressor discharge air enters the blade through flow controlorifices in the base of the blade, passes up through a leading edgepassage, down a midchord passage, up a trailing edge passage, and outthrough holes in the trailing edge of the blade.

The air flow path to the rotor includes an annular passage between aninner liner of the annular combustor of the engine and a gas producerbearing housing, a pre-swirler immediately downstream of the first stagenozzle diaphragm, and a cavity or plenum in front of the rotor.Approximately 60 percent of the air supplied to the cavity is taken onboard the rotor to cool the first stage blades in the manner discussedabove.

The remainder of the air discharged into the cavity in front of therotor flows through apertures in the first stage rotor disc to thedownstream side of the rotor. This air assists in cooling the rotor discand the diaphragm extending between the latter and the second stagenozzles.

The air flowing to and in the cavity on the upstream side of the firststage rotor also operates to cool the first stage rotor disc and thefirst stage nozzle diaphragm of the gas producer turbine.

One advantage of our invention just described is that only one source ofcompressor air is required even though the air is supplied at twodifferent pressures. Also, all flow passages for the cooling air areinternal, i.e., within the engine housings.

One typical conventional arrangment is shown in U.S. Pat. No. 4,023,919issued May 17, 1977, to Patterson; in it external lines are employed tosupply compressor air to the turbine. Furthermore, a different line isrequired to make air available at each of the needed pressures. Systemsof this type are much more complex than ours and much more subject todamage and leakage, all disadvantages of significant and self-evidentimport.

From the foregoing it will be apparent to the reader that one importantand primary object of the invention resides in the provision of novel,improved systems for cooling gas turbine nozzles and rotors and turbinecomponents associated therewith.

Another important, primary, and related object of the invention is theprovision of novel, improved systems for supplying compressor dischargeair to turbine nozzles and rotors of a gas turbine engine and associatedcomponents to cool the nozzles, rotor, and other components.

Still another important, primary object of the invention is theprovision of novel, improved air supply and cooling systems in accordwith the preceding objects which permit high turbine inlet gastemperatures to be used to obtain high thermal efficiencies yet keepturbine component temperatures low enough to give long useful servicelives.

A further, important, primary object of the invention is the provisionof novel, improved air supply and cooling systems in accord with thepreceding object which are particularly adapted to two-shaft, gasturbine engines.

A still further and related, important object of the invention residesin the provision of novel, improved systems which provide optimizedcooling of first and second stage nozzles of a two-shaft gas turbineengine and of the first stage of the rotor of the gas producer turbine.

An important object related to that identified in the precedingparagraph is the provision of novel, improved systems which also provideefficient cooling of the nozzle case of the turbine as well as efficientcooling of diaphragms extending between the first and second nozzlestages and the first stage of the turbine rotor.

Yet another important object of our invention resides in the provisionof novel, improved systems which are simple and relativelynon-susceptible to damage and leakage.

Other important objects and features and additional advantages of thepresent invention will become apparent from the appended claims and asthe ensuing detailed description and discussion proceeds in conjunctionwith the accompanying drawing, in which:

FIGS. 1A and 1B, taken together, constitute a partly sectioned side viewof a gas turbine engine having cooling systems embodying the principlesof the present invention;

FIG. 2 is a fragment of the foregoing view to an enlarged scale;

FIG. 3 is a perspective of an internally cooled first stage nozzle andinsert assembly employed in the engine of FIGS. 1A and 1B;

FIG. 4 is a section through a nozzle in the assembly of FIG. 3 takensubstantially along line 4--4 of the latter Figure;

FIG. 5 is a view similar to FIG. 3 of an internally cooled second stagenozzle and insert assembly employed in the engine of FIGS. 1A and 1B;

FIG. 6 is a side view of the assembly of FIG. 5;

FIG. 7 is a section through a nozzle in the assembly of FIGS. 5 and 6taken substantially along line 7--7 of FIG. 6;

FIG. 8 is a vertical section through a first stage blade of a gasproducer turbine rotor incorporated in the engine of FIGS. 1A and 1B;and

FIG. 9 is a section through the blade taken substantially along line9--9 of FIG. 8.

Referring now the drawing, FIGS. 1A and 1B depict a two-shaft, gasturbine engine 10 equipped with turbine case and nozzle and turbinerotor cooling systems 12 and 14 constructed in accord with and embodyingthe principles of the present invention (see FIG. 2).

Engine 10 includes a fifteen-stage axial flow compressor 16 with aradial-axial inlet 18, inlet guide vanes 19, stators 20, and afifteen-stage rotor 21. The inlet guide vanes 19 and stators 20 aresupported from the compressor housing 22 with the guide vanes andstators 20-1 through 20-5 of the first five stages being pivotallymounted so that they can be adjusted to control the flow of air throughthe compressor.

Each of the fifteen stages of the rotor 21 consists of a disc 23 withradially extending blades 24 fixed to the periphery of the disc. Thestages are integrated into a unitary structure as by electron beamwelding.

The compressor housing is split longitudinally along a vertical planethrough the axial centerline of engine 10 into sections 22a (only one ofwhich is shown) to accommodate assembly of the compressor and tofacilitate inspection, cleaning, and replacement of guide vanes 19 andstators 20 and the blades 24 of compressor rotor 21.

The high pressure air discharged from compressor 16 flows through adiverging diffuser 25 and an enlarged dump plenum 26 to an annularcombustor 27 supported in an insulated combustor case 28.

The compressor discharge air heated by combustor 27 and the combustionproducts generated in the combustor are expanded through a two-stage gasproducer turbine 32 and then through a two-stage power turbine 34. Theturbines are rotatably supported in a nozzle case 36 mounted in anannular turbine housing 38.

As best shown in FIG. 2, the gas producer turbine 32 has a two-stagerotor 42 and stationary, internally cooled, first and second stagenozzles 44 and 46.

The first stage nozzles are integral components of nozzle segments 52each having two nozzles 44. The second stage nozzles are similarlyintegral components of nozzle segments 54 each having two second stagenozzles 46.

The first stage 56 of gas producer turbine rotor 42 includes a disc 58to which internally cooled, radially extending blades 60 are fixed. Anoptical pyrometer 61 is sighted on blades 60 to measure theirtemperature.

The second stage 62 of the rotor includes a disc 63 with uncooled,radially extending blades 64 mounted on its periphery.

The two stages of the gas producer turbine rotor 42 are bolted to eachother (see FIG. 2) and, in cantilever fashion, to the rear end of aforwardly extending shaft 66. Shaft 66 is coupled through rearcompressor hub 68 to compressor rotor 21, thereby drive-connecting gasproducer turbine 32 to the compressor.

The compressor and gas producer turbine are rotatably supported by athrust bearing 70 and by tapered land bearings 72, 74, and 76. Bearings70 and 72 engage the front compressor hub 77 which is bolted to rotor 21and is drive-connected to an accessory drive 78.

Power turbine 34 includes first and second stage nozzles 79 and 80 alsosupported from nozzle case 36 and a rotor 82 having a first bladed stage84 and a second, bladed stage 86 bolted together for concomitantrotation. Rotor 82 is bolted to a power turbine shaft assembly 88rotatably supported by tapered land bearings 90 and 92 and a thrustbearing 94. The shaft assembly is connected through a coupling 96 to anoutput shaft assembly 98 which furnishes the input for a generator,booster compressor, mechanical drive, or other driven unit.

The final major component of turbine engine 10 is an exhaust duct 100for the gases discharged from power turbine 34.

For the most part, the details of the gas turbine engine 10 describedabove are not relevant to the practice of the present invention.Therefore, they will be described only as necessary to provide a settingfor and facilitate an understanding of the latter.

Referring now primarily to FIG. 2, but also to FIGS. 1A and 1B,combustor 27 includes a flame tube 102 surrounded on its outer side byan insulated, annular, outer air liner 104 spaced inwardly from annularcombustor case 28 to form an annular passage 105. Air discharged fromcompressor 16 can bleed from dump plenum 26 through passage 105 to anannular plenum 106 between turbine housing 38 and turbine nozzle case36. A screen assembly 108 composed of a screen 110 supported from thenozzle case by brackets 112 and 114 and a clamp ring 116 isolates plenum106 from the interiors of first and second stage nozzles 44 and 46.

As shown by arrow 118 in FIG. 2, one stream of the compressor dischargeair flows from plenum 106 through transfer tubes 120 into the internallycooled first stage nozzles 44 of the gas producer turbine. A secondstream of the compressor discharge air (see arrows 122 and 124) flowsthrough apertures 126 in an impingement shroud 128 surrounding thenozzle case and impinges against the latter to cool it. As shown byarrow 129, this air then flows through transfer tubes 130 into theinternally cooled, second stage nozzles 46 of the gas producer turbine.

Transfer tubes 120 each include a retainer 131 extending through andfixed to shroud 128 and an air tube 132 extending through nozzle case 36and biased against the outer end of the associated first stage nozzle bya spring 133. Transfer tubes 130 have a dogbone configuration as shownin FIG. 5.

As suggested briefly above and shown in FIG. 3, first stage nozzles 44are incorporated in nozzle segments 52, each of which include twonozzles, an outer shroud 134, and an inner shroud 136. The outer shroudsare seated in recesses in nozzle case 36 with seal strips 138 spanningthe gaps between adjacent segments. The inner shrouds are seated inrecesses in a sliding ring 140 which co-operates with a rim sealassembly 142 and a first stage nozzle diaphragm assembly 144 to isolatethe gas producer turbine from the combustor section of gas turbineengine 10.

As shown in FIG. 4, nozzles 44 are hollow; and each nozzle has an insert146 extending longitudinally therethrough and spanning approximately theleading one-half of the nozzle. The outer ends of the insertscommunicate with the inner or outlet ends of transfer tubes 120,compressor discharge air supplied through the transfer tubes thereforeflowing into inserts 146.

Part of the air thus supplied to each insert exits through openings 148in the leading edges of the insert. As shown by arrows 150, this airflows around the insert and then toward the trailing edge of theassociated nozzle, cooling the leading edge portions of the nozzle. Thisair exits through slots 152 and 154 on the suction and pressure sides ofthe nozzle as shown by arrows 156 and 158 and then flows over theexternal nozzle surfaces toward the trailing edge of the nozzle, coolingthe external surfaces by film cooling.

Pins 160 extend from the side walls of each nozzle 44 into contact withthe insert 146 therein. These pins cause turbulence in the air flowingthrough the nozzles in the manner just described, increasing the rate ofheat transfer between the air and the side walls of the nozzles.

The remainder of the compressor discharge air introduced into the insert146 in each nozzle 44 exits through openings 162 in the trailing edge ofthe insert as shown by arrow 164 in FIG. 4. This air flows toward thetrailing edge of the nozzle, cooling the aft part thereof with pins 160again being employed to promote the rate of heat transfer. This air isdischarged through openings 166 in the trailing edge of the nozzle asshown by arrow 168.

Turning now to FIGS. 5-7, the nozzle segments 54 in which the secondstage nozzles 46 are incorporated also have an outer shroud 170 in whicha wedge-shaped plenum 172 is formed and an inner shroud 174. The outershrouds are fitted in a slot 176 in nozzle case 36, and inner shrouds174 are fitted in recesses formed in an annular, second stage nozzlediaphragm 178. The second stage nozzle diaphragm co-operates with seals180 and 182 bolted to the downstream side of gas producer turbine rotorfirst stage disc 58 to confine the working fluid to a path through thenozzles and turbine blades.

The transfer tubes 130 into which air flows after impinging on casing 36extend through the nozzle case and the front walls 183 of the secondstage turbine segment outer shrouds 170 into fluid communication withthe wedge-shaped inlet plenums 172 formed in the outer shrouds.

From the inlet plenums, the air flows into inserts 184 extendinglongitudinally through nozzles 46. As in the case of nozzles 44, part ofthis air is discharged through openings 185 in the leadig edges of theinserts and flows back along the inserts toward the trailing edges ofthe nozzles as shown by arrows 186 to cool the forward parts of thenozzles. Ribs 187 and 188 spaced longitudinally in the nozzles promoteheat transfer between the air and the nozzle walls.

The remainder of the air is discharged through openings 189 in thetrailing edges of the inserts as shown by arrow 190. This air flowstoward the trailing edges of the nozzles across heat transfer promotingcrossover ribs 192 and, together with the air discharged through theforward ends of the inserts, exits from the nozzles through slots 194 onthe pressure sides of and adjacent the trailing edges of the nozzles asshown by arrow 196.

Referring again to FIG. 2, openings are also formed through the innershrouds 174 of second stage nozzle segments 54. Part of the air suppliedto the inserts 184 in the nozzles 46 of these segments is dischargedthrough these openings as shown by arrow 198 to assist in cooling secondstage nozzle diaphragm 178.

Also, as shown in FIG. 2, a part of the air flowing through impingementshroud 128 to cool nozzle case 36 flows through annular gaps 200 innozzle case 36 around transfer tubes 120 as shown by arrow 202. This airexits through openings 204 in internal flow barrier defining, outershroud rear walls 206 (see FIG. 3) and flows across the inner surfacesof tip shoe segments 207 surrounding gas producer turbine rotor firststage 52 to cool the tip shoe segments.

The gaps between the tip shoes and, also, the gaps between the secondstage nozzle segments 54 are closed by seal strips like those identifiedby reference character 138 in FIG. 2.

Referring again to FIGS. 1A and 1B and FIG. 2, an inner air liner 208, aring 209, and covers 210 and 211 form an inner, annular, passage 212through which compressor discharge air can flow from dump plenum 26 tofirst stage nozzle diaphragm 144. This air is employed in the coolingsystem 14 for the first stage 56 of gas producer turbine rotor 42.

From plenum 212 the air flows through apertures 214 in diaphragm 144 andswirlers 216 into a plenum or cavity 218 defined by the diaphragm, rimseal assembly 142, the disc 58 of the rotor first stage 52, andlabyrinth seals 220 and 222 bolted to the upstream side of the latter.Swirlers 216 are employed to impart a tangential component to the airflowing through apertures 214 so that, when it is taken on board thefirst stage of the gas producer turbine rotor, its velocity componentwill approximate that of the rotor, minimizing viscous drag on thelatter.

The air in plenum 218 assists in cooling first stage nozzle diaphragm144 and gas producer turbine first stage rotor disc 58 and flowsupwardly through passages 224 in seal 220 as shown by arrow 226 to theroot or inner ends of first stage, internally cooled, rotor blades 60.

Referring now to FIGS. 8 and 9, blades 60 have internal cavities 227,228, and 230 in which inserts or trip strips 232, 234, and 236 aremounted. Air flowing radially outward through passages 224 is introducedinto the leading edge cavities 227 in blades 60 through inlet passagesor orifices 238 as shown by arrow 240. This air flows outwardly throughcavity 227, reverses its direction as shown by arrow 242, flows inwardlythrough mid-chord cavity 228 as shown by arrow 244, then again reversesdirection as shown by arrow 245, and flows outwardly through slots 246in the trailing edge 247 of the blade as shown by arrow 248.

A part of the air introduced into passages 224 is diverted throughlongitudinally extending passages 254 in disc 58 and radial passages 255through seal 180 into a plenum 256 defined by the disc, seals 180 and182, and second stage nozzle diaphragm 178. This air is employed to coolthe disc, to assist in cooling the diaphragm, and for hot gas bufferingbetween the disc and diaphragm. As indicated above, about 40 percent ofthe air flowing to the first stage rotor will be employed in this mannerin the exemplary gas turbine engine 10 described above and illustratedin FIG. 1.

The invention may be embodied in other specific forms without departingfrom the spirit or essential characteristics thereof. The presentembodiment is therefore to be considered in all respects as illustrativeand not restrictive, the scope of the invention being indicated by theappended claims rather than by the foregoing description; and allchanges which come within the meaning and range of equivalency of theclaims are therefore intended to be embraced therein.

What is claimed and desired to be secured by Letters Patent is:
 1. A gasturbine comprising: a nozzle case through which heated gases are adaptedto flow to drive a turbine rotor rotatably supported in said case; ahousing surrounding said case in spaced relation thereto; first andsecond stage, flow directing nozzles supported in longitudinally spacedrelationship from said case; and means for cooling said nozzles and saidcase comprising means for directing a first stream of air directly fromthe space between said housing and said case to the interiors of thenozzles in one of said stages, means in the space between said housingand said case and surrounding the latter for directing a second,separate stream of air from said space into impinging relationship withsaid case in the span between said nozzle stages around the periphery ofsaid case, and means for then directing said second stream of air intothe interiors of the nozzles in the other of said stages.
 2. A gasturbine as defined in claim 1 in which the means for directing saidsecond stream of air into impinging relationship with said casecomprises an apertured shroud spaced from said case over the aforesaidspan and wherein the means for directing air into the nozzles in theother of said stages has an inlet communicating with the space betweensaid case and said apertured shroud.
 3. A gas turbine as defined inclaim 2 wherein the means for directing air to the interiors of saidnozzles in said one of said stages has inlets communicating with thespace between the nozzle case and the housing at intervals therearound,said turbine further comprising a screen so surrounding said inlets andsaid apertured shroud in the space between the case and the housing asto intercept air flowing from said space into said inlets through saidshroud and thereby keep foreign material from reaching and plugging theapertures in the shroud or the nozzles of the first and second stages.4. A gas turbine comprising: a nozzle case through which heated gasesare adapted to flow to drive a turbine rotor rotatably supported in saidcase; a housing surrounding said case in spaced relation thereto; anannular array of flow directing nozzles supported in and from said case;and means for cooling said nozzles and at least one other turbinecomponent housed in said nozzle case, said nozzles having portions whichco-operate with said nozzle case to form an internal flow barrier in thenozzle case and said cooling means comprising transfer tubes, the inletsof said transfer tubes being in fluid communication with the spacebetween said case and said housing and the outlets of said tubes beingin fluid communication with the interiors of said nozzles, apertures insaid nozzle portions via which air can flow through said internalbarrier into the interior of said nozzle case and into convective heattransfer relationship with said other turbine component therein, andmeans for supplying air from the space between said housing and saidnozzle case through the latter to said apertures and for keeping theheated gases in said nozzle case from escaping through said apertures,the last-mentioned means comprising annular passages through said nozzlecase around said transfer tubes.
 5. A gas turbine as defined in claim 4which includes a second annular array of nozzles supported in and fromsaid nozzle case in spaced relation to the aforesaid array of nozzlesand means for cooling the nozzles in said second array comprising meansfor directing air from the space between the nozzle case and the housinginto said nozzles, the nozzles in the second array thereof having plenumchambers at the outer ends thereof and the means for directing air intosaid nozzles comprising transfer tubes extending in gastightrelationship therewith through said nozzle case, the inlet ends of saidtransfer tubes being in fluid communication with the space between saidcase and said housing and the outlet ends thereof being in fluidcommunication with said plenum chambers.
 6. A gas turbine comprising: anozzle case through which heated gases are adapted to flow to drive aturbine rotor rotatably supported in said case; a housing surroundingsaid case in spaced relation thereto; an annular array of flow directingnozzles supported in and from said case; and means for cooling saidnozzles comprising means for directing a stream of air from the spacebetween said housing and said case to the interiors of the nozzles, saidnozzles being hollow and said turbine further comprising a single insertdisposed in and extending longitudinally of each of said nozzles, saidinserts being in fluid communication at the outer ends thereof with themeans for directing air into the nozzles in which said inserts aredisposed; and there being apertures in the leading edges of said insertsthrough which air can flow to cool the leading edge portions of saidnozzles; slots in the pressure and suction sides of said nozzles betweenthe leading and trailing edges thereof adapted to discharge said airalong and into heat transfer relationship with the exterior surfaces ofsaid nozzles; openings in the trailing edges of said inserts throughwhich the remainder of the air introduced into the inserts can exittherefrom into the interiors of the nozzles; and exits for said air inthe trailing edges of the nozzles.
 7. A gas turbine as defined in claim6 which includes a second annular array of nozzles supported in and fromsaid nozzle case in spaced relation to the aforesaid array of nozzlesand means for cooling the nozzles in said second array comprising meansfor directing air from the space between the nozzle case and the housinginto said nozzles, the nozzles in the second array thereof having plenumchambers at the outer ends thereof; there being a single insert in andextending longitudinally of each of said nozzles, said inserts being influid communication with said plenum chambers at the outer ends thereof;and there being openings in the leading edges of said inserts throughwhich air can flow into the nozzles to cool the leading edge portionsthereof, openings in the trailing edges of the inserts through which theremainder of the air introduced into the inserts can flow into thenozzles to cool the trailing edge portions thereof, and dischargeopenings in the nozzles adjacent the trailing edges thereof for the airintroduced into the nozzles.
 8. A gas turbine comprising: a nozzle casethrough which heated gases are adapted to flow; a housing surroundingsaid case in spaced relation thereto; first stage, flow directingnozzles integrated into nozzle segments supported from said case; abladed rotor rotatably supported in said nozzle case adjacent said firststage nozzles; tip shoe means supported from said case in surroundingrelationship to said rotor; means for cooling said nozzles comprisingmeans for directing air from the space between said housing and saidcase into the interiors of the nozzles, said means comprising transfertubes disposed in passages extending through the nozzle case andcommunicating at opposite ends thereof with the space between the caseand the housing and with the nozzle interiors; and means for coolingsaid tip shoes comprising means for diverting air through said nozzlecase around said transfer tubes and apertures in the trailing edges ofsaid segments for directing said diverted air across the exposedsurfaces of said tip shoe means.
 9. A gas turbine as defined in claim 8together with means for directing said diverted air into heat transferrelationship with the outer side of said nozzle case before said airpasses through the apertures in the trailing edges of the nozzlesegments.
 10. A gas turbine as defined in claim 8 together with meansfor reducing the pressure on the diverted air before said air isdiverted around said transfer tubes and through the apertures in thetrailing edges of the nozzle segments.
 11. A gas turbine comprising: anozzle case through which heated gases are adapted to flow; a turbinerotor comprising at least one stage which includes a disc and an annulararray of hollow blades fixed to and extending radially from said disc,said rotor being rotatably supported in said nozzle case and adapted tobe driven by said heated gases; a housing surrounding said nozzle casein spaced relation thereto; first and second stage, flow directingnozzles respectively supported from said case upstream of the downstreamfrom said bladed rotor stage; an annular diaphragm in sealing engagementwith said second stage nozzles and the downstream side of said disc forconfining said heated gases to a path between said rotor blades and saidsecond stage nozzles; means for cooling said nozzles and said nozzlecase comprising means for conducting a first stream of air from thespace between said housing and said nozzle case to the interiors of thenozzles in one of said nozzle stages and means for directing a secondstream of air from said space into impinging relationship with said caseand then into the interiors of the nozzles in the other of said nozzlestages; and means for cooling said one rotor stage and said diaphragmcomprising means for directing part of a third stream of cooling airinto said hollow blades and means for directing the remainder of saidthird stream of air serially through said disc and into convective heattransfer relationship with the downstream side of said disc and saiddiaphragm.
 12. A gas turbine comprising: a nozzle case through whichheated gases are adapted to flow; a turbine rotor having bladed firstand second stages, said rotor being rotatably supported in said nozzlecase and adapted to be driven by said gases; a housing surrounding saidnozzle case in spaced relation thereto; first and second stage, flowdirecting nozzles supported from said case adjacent and upstream fromthe first and second rotor stages, respectively; means for cooling saidnozzles and said case comprising means for conducting a first stream ofair from the space between said housing and said case to the interiorsof the nozzles in one of said stages and means for directing a secondstream of air from said space into impinging relationship with said caseand then into the interiors of the nozzles in the other of said stages;a first stage nozzle diaphragm for isolating the upstream end of saidturbine; and means for cooling the first stage of the turbine rotorcomprising means for directing a third stream of air through saiddiaphragm and into the interiors of the blades of said first stagethrough the root ends thereof, said first rotor stage also comprising adisc to which the blades of said stage are attached; said turbine alsohaving a second stage nozzle diaphragm adjacent the downstream side ofand co-operating with the first rotor stage and the second stage nozzlesto keep gases exiting from the first rotor stage from bypassing thesecond stage nozzles; and the means for directing air to the blades ofthe first rotor stage also including means for effecting a flow of partof said air through the first rotor stage disc into the space betweensaid disc and said diaphragm and into convective heat transferrelationship therewith to cool the aforesaid components.
 13. A gasturbine comprising: a nozzle case through which heated gases are adaptedto flow; a turbine rotor stage rotatably supported in said case andadapted to be driven by said heated gases; a tip shoe surrounding saidrotor; a housing surrounding said case in spaced relation thereto; firstand second stage, flow directing nozzles supported in longitudinallyspaced relationship from said case on the upstream and downstream sidesof said turbine rotor stage, respectively; and means for cooling saidnozzles, said tip shoes, and said case comprising means for directing afirst stream of air from the space between said housing and said case tothe interiors of nozzles in one of said stages, means for directing asecond stream of air from said space into impinging relationship withsaid case, means for then directing a part of said second stream of airinto the interiors of the nozzles in the other of said stages, and meansfor directing the rest of said second stream of air into heat transferrelationship with said tip shoe.